CN114060170A - Open type staged combustion air-extraction circulation liquid rocket engine - Google Patents
Open type staged combustion air-extraction circulation liquid rocket engine Download PDFInfo
- Publication number
- CN114060170A CN114060170A CN202111251559.8A CN202111251559A CN114060170A CN 114060170 A CN114060170 A CN 114060170A CN 202111251559 A CN202111251559 A CN 202111251559A CN 114060170 A CN114060170 A CN 114060170A
- Authority
- CN
- China
- Prior art keywords
- thrust chamber
- gas
- oxidant
- fuel
- turbine
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/46—Feeding propellants using pumps
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/46—Feeding propellants using pumps
- F02K9/48—Feeding propellants using pumps driven by a gas turbine fed by propellant combustion gases or fed by vaporized propellants or other gases
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/44—Feeding propellants
- F02K9/56—Control
- F02K9/58—Propellant feed valves
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/60—Constructional parts; Details not otherwise provided for
- F02K9/62—Combustion or thrust chambers
- F02K9/64—Combustion or thrust chambers having cooling arrangements
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02E—REDUCTION OF GREENHOUSE GAS [GHG] EMISSIONS, RELATED TO ENERGY GENERATION, TRANSMISSION OR DISTRIBUTION
- Y02E20/00—Combustion technologies with mitigation potential
- Y02E20/34—Indirect CO2mitigation, i.e. by acting on non CO2directly related matters of the process, e.g. pre-heating or heat recovery
Abstract
The invention relates to an open type staged combustion air extraction circulation liquid rocket engine, which comprises a set of turbine pump, a staged combustion thrust chamber, corresponding control valves and the like, wherein a fuel gas generator is not arranged, and a turbine of the turbine pump is driven by fuel-rich gas directly led out from an upper cavity of the thrust chamber. The staged combustion thrust chamber adopts a two-stage combustion mode of an upper chamber and a lower chamber, part of oxidant and all fuel are subjected to rich combustion in the upper chamber of the thrust chamber to form a rich combustion area with relatively low temperature, and the generated rich combustion gas is introduced into the lower chamber of the thrust chamber and is subjected to afterburning combustion with most of the oxidant introduced from the back of the oxygen pump. The mixing ratio of the cavities in the thrust chamber is selected according to the temperature of the rich combustion gas which can be borne by the turbine. The present invention has the features of simple system, less components, no fuel gas generator component, gas-liquid combustion in the main combustion area of the thrust chamber, high combustion stability, etc. and can lower the cost of engine greatly and raise the thrust-weight ratio and reliability of engine.
Description
Technical Field
The invention belongs to a rocket engine, and relates to an open type staged combustion air extraction circulation liquid rocket engine which can be used for a liquid propellant pumping type rocket engine of a commercial carrier rocket and a space vehicle.
Background
The system circulation modes of the pumping pressure type engine adopted by the current carrier rocket and spacecraft mainly comprise open circulation and closed circulation. The closed cycle has high performance but complex system, and the open cycle liquid rocket engine has relatively low performance although the turbine exhaust loss exists, but the system is simple, the manufacturing cost is low, and the thrust-weight ratio can reach a relatively high level, so that the closed cycle liquid rocket engine still has a main choice for commercial aerospace at present, for example, the open cycle mode is adopted by the Meilin engine used by the falcon Jiu commercial carrier rocket of the space X company in the United states. There are two main ways of an open cycle liquid rocket engine: one is the gasifier cycle and the other is the extraction cycle.
The gas generator cycle generally consists of an independent gas generator, a turbopump, a thrust chamber, a valve and the like, for example, the conventional propellant rocket engine YF-20 still used in China; and the gas extraction circulation has no gas generator, so the system is simpler. The gas-extraction circulation engine can drive the turbine by extracting gas from the main combustion chamber, but because the temperature of the gas in the main combustion chamber is very high and generally exceeds 3000K, the gas with ultrahigh temperature cannot be used for driving the turbine in engineering, so the gas-extraction circulation has no application example internationally. The low-temperature propellant medium is used for heating, gasifying and expanding a cooling jacket of a thrust chamber to drive a turbine (the gaseous propellant medium after the turbine is driven is directly discharged to the outside) and is also an air extraction circulation mode without a fuel gas generator, but the air extraction circulation engine does not lead out fuel gas from a main combustion chamber, the propellant medium for driving the turbine does not participate in combustion, the working capacity of the engine depends on the expansion working capacity of the gaseous propellant, and the engine is only suitable for engines with low-temperature propellant media, such as liquid oxygen/liquid hydrogen and liquid oxygen/methane propellant combinations. One such type of suction cycle engine is the LE-9 open expansion cycle engine of Japan.
The invention provides an air extraction circulation engine without an independent fuel gas generator, which leads low-temperature fuel gas out of a main combustion chamber by organizing the main combustion chamber to carry out staged combustion. The novel air extraction circulation engine has the characteristics of simple and reliable system, wide adaptability and the like, and is suitable for the combination of double low-temperature propellants and the combination of liquid oxygen/kerosene propellants and normal-temperature propellants.
Disclosure of Invention
Technical problem to be solved
In order to avoid the defects of the prior art, the invention provides an open type staged combustion air extraction circulation liquid rocket engine, aiming at greatly simplifying a liquid rocket engine system and solving the problems of relatively high complexity and high reliability guarantee difficulty of the engine system for a carrier rocket.
Technical scheme
An open type staged combustion air extraction circulation liquid rocket engine is characterized by comprising a staged combustion thrust chamber 6, an igniter 3, a turbopump and starter 1, relevant valves and pipelines; the staged combustion thrust chamber 6 adopts a two-stage combustion mode and sequentially comprises a thrust chamber upper chamber 6-1, a gas-liquid injector 6-2, a thrust chamber lower chamber 6-3 and a tail nozzle 6-4 from top to bottom, wherein part of oxidant and all fuel are organized to be burnt rich in fuel in the thrust chamber upper chamber 6-1, the gas-liquid injector 6-2 is provided with a fuel rich gas injection unit and an oxidant injection unit, the fuel rich gas generated in the thrust chamber upper chamber 6-1 and most of oxidant after an oxidant pump enter the thrust chamber lower chamber 6-3 through the gas-liquid injector 6-2 to be subjected to afterburning combustion, the generated gas is accelerated and expanded through the tail nozzle and then is discharged out of the thrust chamber, the periphery of the thrust chamber is provided with a thrust chamber fuel cooling channel 6-5, and an outlet is communicated with the thrust chamber upper chamber 6-1; the external fuel is connected with the inlet of the thrust chamber fuel cooling channel 6-5 through a fuel pump 8 and a fuel valve 9; the external oxidant is respectively communicated with the upper thrust chamber cavity 6-1 and the gas-liquid injector 6-2 through an oxidant pump 7, an oxygen auxiliary valve 4 and an oxygen main valve 5; the turbine pump comprises a turbine 2, a fuel pump 8 and an oxidant pump 7 which are coaxially connected in series with the turbine, and a starter 1 is connected with a collector of the turbine; the turbine of the turbine pump is driven by rich fuel gas directly led out from the upper cavity 6-1 of the staged combustion thrust chamber, and the rich fuel gas is directly discharged to the outside after driving the turbine.
The starter comprises a gunpowder starter or a high-pressure gas cylinder starter.
The outer part of the thrust chamber fuel cooling channel 6-5 is provided with a thrust chamber oxidant cooling channel 6-6, and the outer oxidant is communicated with the gas-liquid injector 6-2 through an oxidant pump 7, an oxygen main valve 5 and the thrust chamber oxidant cooling channel 6-6.
The external oxidant is communicated with a main valve 12 through an oxidant pump 7, is directly communicated with the upper cavity 6-1 of the thrust chamber, and is communicated with a gas-liquid injector 6-2 through an oxygen main valve 5.
Advantageous effects
The invention provides an open type staged combustion air extraction circulation liquid rocket engine which comprises a set of turbine pump, a staged combustion thrust chamber, corresponding control valves and the like, wherein a fuel gas generator is not arranged, and a turbine of the turbine pump is driven by fuel-rich gas directly led out from an upper cavity of the thrust chamber. The staged combustion thrust chamber adopts a two-stage combustion mode of an upper chamber and a lower chamber, namely, part of oxidant and all fuel are subjected to rich combustion in the upper chamber of the thrust chamber to form a rich combustion region with relatively low temperature, and the generated rich combustion gas is introduced into the lower chamber of the thrust chamber and is subjected to afterburning combustion with most of the oxidant introduced from the back of the oxygen pump. The mixing ratio of the cavities in the thrust chamber is selected according to the temperature of the rich combustion gas which can be borne by the turbine. The propellant adopted by the open type staged combustion air extraction circulation liquid rocket engine can be liquid oxygen/liquid hydrogen, liquid oxygen/methane, liquid oxygen/kerosene, dinitrogen tetroxide/unsymmetrical dimethylhydrazine or other propellant combinations.
Has the advantages that: the open type staged combustion air extraction circulation liquid rocket engine provided by the invention has the characteristics of simple system, few components, no fuel gas generator component, gas-liquid combustion in a main combustion area of a thrust chamber, good combustion stability and the like, and can greatly reduce the engine cost and improve the thrust-weight ratio and the reliability of the engine.
Compared with the conventional open cycle engine-gas generator cycle, the open staged combustion gas extraction cycle liquid rocket engine provided by the invention can reduce one gas generator, a gas generator fuel and oxidant supply valve and a gas generator ignition device (for non-self-ignition propellant combination), thereby further reducing the complexity of an engine system, and being more beneficial to reducing the cost and improving the thrust-weight ratio and the reliability of the engine. Compared with the open type air extraction expansion cycle, the open type staged combustion air extraction cycle liquid rocket engine provided by the invention has wide adaptability, is suitable for low-temperature propellant combination and normal-temperature propellant combination, and has stronger expansion working capacity and relatively less energy loss because the driving turbine medium participates in partial combustion for double-low-temperature propellant combination, thereby being more beneficial to improving the performance of the engine.
Drawings
FIG. 1: scheme-system schematic diagram of open type staged combustion air extraction circulation liquid rocket engine
FIG. 2: scheme two-system schematic diagram of open type staged combustion air extraction circulation liquid rocket engine
FIG. 3: scheme three-system schematic diagram of open type staged combustion air extraction circulation liquid rocket engine
FIG. 4: is a four-system schematic diagram of a scheme of an open type staged combustion air extraction circulation liquid rocket engine
Wherein: 1. a gunpowder starter; 2. a turbine; 3. a spark plug igniter; 4. an oxygen secondary valve; 5. an oxygen valve; 6. a staged combustion thrust chamber; 6-1, an upper thrust chamber cavity; 6-2, a gas-liquid injector; 6-3, a lower cavity of the thrust chamber; 6-4. a tail nozzle; 6-5, a thrust chamber fuel cooling channel; 6-6, a thrust chamber oxidant cooling channel; 7. an oxidant pump; 8. a fuel pump; 9. a fuel valve; 10. a high-pressure nitrogen cylinder; 11. a cylinder valve; 12. an oxygen main valve.
Detailed Description
The invention will now be further described with reference to the following examples and drawings:
the invention is shown in figure 1. In the figure, a liquid oxygen/kerosene propellant combination is taken as an example, and a propellant combination such as hydrogen peroxide/hydrocarbons, nitrogen oxide/hydrocarbons, liquid oxygen/liquid hydrogen, liquid oxygen/methane, and the like may be used. The present invention mainly consists of oxidant pump, fuel pump, turbine, staged combustion thrust chamber, oxygen main valve, oxygen auxiliary valve, fuel valve, spark plug igniter, powder starter and other components.
The external fuel storage tank is connected with the inlet of the fuel pump through a front pipeline of the fuel pump, the fuel pump drives fuel to flow, the outlet of the fuel pump is connected with the inlet of the fuel cooling channel of the thrust chamber through a rear pipeline of the fuel pump, and a fuel valve is arranged in the rear pipeline of the fuel pump and used for controlling the circulation of the fuel channel.
The external oxidant storage tank is connected with an oxidant pump inlet through an oxidant pump front pipeline, the oxidant pump drives oxidant to flow, an oxidant pump outlet is respectively connected with the thrust chamber upper cavity and the gas-liquid injector through a pump rear pipeline, an oxygen auxiliary valve is arranged in the pump rear pipeline connected with the thrust chamber upper cavity and used for controlling the circulation of the oxygen auxiliary channel, and an oxygen main valve is arranged in the pump rear pipeline connected with the gas-liquid injector and used for controlling the circulation of the oxygen main channel.
The staged combustion thrust chamber comprises a thrust chamber upper cavity, a gas-liquid injector, a thrust chamber lower cavity and a tail nozzle from top to bottom in sequence, propellant in the thrust chamber upper cavity is subjected to rich combustion, the gas-liquid injector is used for blocking the thrust chamber upper cavity and the thrust chamber lower cavity and simultaneously injecting main path oxidant and rich combustion gas, and the propellant in the thrust chamber lower cavity is subjected to afterburning combustion. In addition, the thrust chamber fuel cooling channel is distributed inside the whole wall surface of the staged combustion thrust chamber and is used for reducing the temperature of the wall surface of the staged combustion thrust chamber. And a spark plug igniter is arranged in the upper thrust chamber cavity and is used for igniting and burning the propellant in the upper thrust chamber cavity.
One side of a collector of the turbine is connected with the upper cavity of the thrust chamber through a pipeline, the turbine is driven by rich combustion gas, the other side of the collector is connected with the gunpowder starter, the turbine is started when the engine is started, and an outlet of the turbine is directly connected with the outside. The turbine, the fuel pump and the oxidant pump are coaxially and sequentially installed to form a turbine pump, and the turbine rotates and simultaneously drives the fuel pump and the oxidant pump to rotate.
The working principle is as follows:
before starting, the fuel pump is filled with fuel before the fuel valve, the oxidant pump is filled with oxidant before the oxidant main valve and the oxygen auxiliary valve, when starting, the gunpowder starter works, the turbine drives the oxidant pump and the fuel pump to start rotating, the pressure after the pump rises, when the pressure after the pump rises to a certain value, the fuel valve, the oxygen auxiliary valve and the oxygen main valve are opened in sequence, and the fuel is filled into the fuel cooling channel of the thrust chamber and then all enters the upper chamber (rich combustion chamber) of the thrust chamber; the oxidant behind the oxidant pump is divided into two paths, one path (a small part) enters an upper thrust chamber cavity (a rich combustion cavity) through an oxygen auxiliary valve to ignite and combust with the fuel to generate rich combustion gas with relatively low temperature, and the other path (a large part) enters a lower thrust chamber cavity (a gas/liquid combustion cavity) through an oxygen main valve to fully combust with the rich combustion gas flowing in from the upper thrust chamber cavity and is sprayed out through a tail spray pipe of the thrust chamber to generate thrust; most of the rich-combustion gas generated on the upper cavity of the thrust chamber enters the downstream to be subjected to gas/liquid combustion with the oxidant, and a small part of the rich-combustion gas is led to the collector of the turbine, drives the turbine through the turbine nozzle, and directly sprays the rich-combustion gas behind the turbine out of the environment. After the pressure of the staged combustion thrust chamber is built up, the pressure of the fuel-rich gas for driving the turbine rises, the gunpowder starter stops working, the turbine is driven by the fuel-rich gas led out from the upper cavity of the thrust chamber alone, and the engine reaches a stable working state. When the engine is shut down, the oxygen main valve, the oxygen auxiliary valve and the fuel valve are closed in sequence, the pressure of the staged combustion thrust chamber is reduced, the pressure of the fuel-rich gas driving the turbine is reduced along with the pressure, and the engine is shut down.
Taking a liquid oxygen/kerosene propellant combination as an example, according to the principle explained in the technical scheme, an engine assembly is configured according to the attached drawings 1 to 4 (a necessary blow-off valve and a blow-off air source system can be added). The medium of the engine driving turbine is extracted from the upper cavity of the thrust chamber, the upper cavity mixing ratio of the thrust chamber is 0.4 in consideration of the upper limit of the gas temperature which can be borne by the turbine, the corresponding rich-combustion gas temperature is 1100K, and the flow of the driving turbine can be controlled by the diameter of the throat part of the turbine nozzle and the mode of adding a throttling ring on the rich-combustion gas extraction pipeline according to the balance calculation result of system parameters. The lower cavity of the thrust chamber is an afterburning area (most of the oxidant introduced after the pump performs afterburning with the rich fuel gas flowing in the upper cavity), and the mixing ratio of the afterburning area and the lower cavity of the thrust chamber can be selected according to the conventional principle of approaching the theoretical mixing ratio of the thrust chamber (which can be 2.6 in the example).
The liquid oxygen kerosene is a non-spontaneous combustion propellant, the ignition of the upper cavity of the thrust chamber can adopt the modes of gunpowder ignition, spark plug torch ignition or chemical ignition and the like, and the lower cavity of the thrust chamber is formed by combusting the liquid oxygen with rich fuel gas with the temperature of 1100K without arranging an igniter.
The oxidizer pump and the fuel pump are single-stage centrifugal pumps, the turbine is an impact turbine, and the starter of the turbine pump can be started by a gunpowder starter, high-pressure helium or high-pressure nitrogen. The fuel valve, the oxygen main valve and the oxygen auxiliary valve can realize the switching function by adopting an electric control pneumatic valve.
The first specific embodiment is as follows:
as shown in the attached figure 1, the invention mainly comprises a gunpowder starter 1, a turbine 2, a spark plug igniter 3, an oxygen auxiliary valve 4, an oxygen main valve 5, a staged combustion thrust chamber 6, an oxidant pump 7, a fuel pump 8, a fuel valve 9 and other components. The staged combustion thrust chamber 6 comprises a thrust chamber upper cavity 6-1, a gas-liquid injector 6-2, a thrust chamber lower cavity 6-3, a tail nozzle 6-4 and a thrust chamber fuel cooling channel 6-5.
The inlet of the fuel pump 8 is connected to an external fuel tank through a pre-pump line of the fuel pump 8, and the fuel pump 8 rotates to raise the fuel pressure and drive the fuel to flow through the line. The outlet of the fuel pump 8 is connected with the inlet of the fuel cooling channel 6-5 of the thrust chamber through a rear pump pipeline, and a fuel valve 9 is arranged in the rear pump pipeline and used for controlling the circulation of the fuel pipeline. The fuel flows through the thrust chamber fuel cooling channels 6-5 for reducing the staged combustion thrust chamber temperature and then all goes directly into the thrust chamber upper chamber 6-1.
An inlet of an oxidant pump 7 is connected with an external oxidant storage tank through a pre-pump pipeline of the oxidant pump 7, the oxidant pump 7 raises the pressure of an oxidant after rotating to drive the oxidant to flow in the pipeline, an outlet of the oxidant pump 7 is connected with a post-pump pipeline, the post-pump pipeline is only one pipeline, and is divided into two paths before being connected to a staged combustion thrust chamber 6, wherein one path is an oxygen main path and is connected with an upper chamber 6-1 of the thrust chamber, the other path is an oxygen auxiliary path and is connected with a gas-liquid injector 6-2, an oxygen auxiliary valve 4 is arranged in a post-pump pipeline connected with the upper chamber 6-1 of the thrust chamber, the circulation of the oxygen auxiliary path is controlled, and an oxygen main valve 5 is arranged in a post-pump pipeline connected with the gas-liquid injector 6-2 to control the circulation of the oxygen main path. And a small part of the oxidant passing through the oxidant pump 7 enters the upper cavity 6-1 of the thrust chamber through the oxygen secondary path, and a large part of the oxidant enters the gas-liquid injector 6-2 and then is injected into the lower cavity 6-3 of the thrust chamber.
The staged combustion thrust chamber 6 comprises a thrust chamber upper cavity 6-1, a gas-liquid injector 6-2, a thrust chamber lower cavity 6-3 and a tail nozzle 6-4 from the head to the tail in sequence, and all the components are coaxially arranged. In the upper cavity 6-1 of the thrust chamber, all fuels and a small part of oxidant are subjected to rich combustion, most of generated rich combustion gas enters the lower cavity 6-3 of the thrust chamber through the gas-liquid injector 6-2 and continuously reacts with the oxidant injected by the gas-liquid injector 6-2, and the rich combustion gas is accelerated and expanded through the tail nozzle 6-4 after complete combustion and then is discharged out of the engine. The gas-liquid injector 6-2 is used for blocking an upper thrust chamber cavity 6-1 and a lower thrust chamber cavity 6-3 and simultaneously comprises an oxidant injection unit and a fuel-rich gas injection unit. The upper thrust chamber cavity 6-1 is provided with a spark plug igniter 3 for igniting and burning the fuel and the oxidant in the upper thrust chamber cavity 6-1.
One side of a collector of the turbine 2 is connected with an upper thrust chamber cavity 6-1 through a pipeline, and a small part of rich gas generated in the upper thrust chamber cavity 6-1 enters the collector of the turbine 2 to drive the turbine. The other side of the collector of the turbine 2 is connected with the gunpowder starter 1, when the turbine 2 is started, the gunpowder starter 1 is ignited, the generated gas enters the collector of the turbine 2 and then starts the turbine, the outlet of the turbine 2 is directly connected with the outside, and all gas passing through the turbine 2 is discharged to the outside. The turbine 2, the fuel pump 8 and the oxidizer pump 7 are coaxially and sequentially installed and assembled to form a turbine pump system, and a sealing device is arranged among the three. When the engine works, the turbine 2 rotates and drives the fuel pump 8 and the oxidant pump 7 to rotate simultaneously, so as to drive fuel and oxidant.
The second specific embodiment:
as shown in fig. 2, the present invention mainly comprises a turbine 2, a spark plug igniter 3, an oxygen auxiliary valve 4, an oxygen main valve 5, a staged combustion thrust chamber 6, an oxidizer pump 7, a fuel pump 8, a fuel valve 9, a high pressure nitrogen gas cylinder 10, a cylinder valve 11, and other components. The difference between the second scheme and the first scheme is that the second scheme is lack of the gunpowder starter 1, but a high-pressure nitrogen gas cylinder 10 and a gas cylinder valve 11 are added, the high-pressure nitrogen gas cylinder 10 is connected with a collector of the turbine 2 through a pipeline, and the gas cylinder valve 11 is arranged in the pipeline. In the second scheme, the starting process of the turbine 2 is changed into: and (3) opening a gas cylinder valve 11, enabling nitrogen in a high-pressure nitrogen cylinder 10 to enter a collector of the turbine 2 through a pipeline and then driving the turbine 2 to move, and starting the turbine 2. The other parts are the same as the first scheme.
The third concrete implementation scheme is as follows:
as shown in figure 3, the invention mainly comprises a gunpowder starter 1, a turbine 2, a spark plug igniter 3, an oxygen auxiliary valve 4, an oxygen main valve 5, a staged combustion thrust chamber 6, an oxidant pump 7, a fuel pump 8, a fuel valve 9 and other components. The staged combustion thrust chamber 6 comprises a thrust chamber upper cavity 6-1, a gas-liquid injector 6-2, a thrust chamber lower cavity 6-3, a tail nozzle 6-4, a thrust chamber fuel cooling channel 6-5 and a thrust chamber oxidant cooling channel 6-6. The difference between the third scheme and the first scheme is that a thrust chamber oxidant cooling channel 6-6 is added in a staged combustion thrust chamber 6 in the third scheme, the thrust chamber oxidant cooling channel 6-6 and the thrust chamber fuel cooling channel 6-5 are two independent cooling channels which are isolated from each other, and the thrust chamber oxidant cooling channel 6-6 is connected with a pipeline behind an oxygen main valve 5. The oxidant of the oxygen main path does not directly enter the gas-liquid injector 6-2 after passing through the oxygen main valve 5, but enters the thrust chamber oxidant cooling channel 6-6 from the rear section or throat of the tail nozzle 6-4, and completely enters the gas-liquid injector 6-2 after flowing through the thrust chamber oxidant cooling channel 6-6. The other parts are the same as the first scheme.
The fourth specific embodiment:
as shown in fig. 4, the present invention is mainly composed of a powder starter 1, a turbine 2, a spark plug igniter 3, an oxygen main valve 5, a staged combustion thrust chamber 6, an oxidizer pump 7, a fuel pump 8, a fuel valve 9, an oxygen main valve 12, and the like. The difference between the scheme four and the scheme one is that the oxygen auxiliary valve 4 is absent in the scheme four, but the oxygen main valve 12 is added, the oxygen main valve 12 is arranged before the pump rear pipeline of the oxidant pump 7 is not divided into two paths, the oxygen auxiliary pipeline is not provided with a valve, and the oxygen main valve 5 is still arranged in the oxygen main pipeline. The oxygen main valve 12 controls the circulation of all the oxidant, after the oxygen main valve 12 is opened, the oxidant can enter the upper cavity 6-1 of the thrust chamber, at the moment, the oxygen main valve 5 is opened, and the oxidant can enter the gas-liquid injector 6-2. The other parts are the same as the first scheme.
Claims (4)
1. An open type staged combustion air extraction circulation liquid rocket engine is characterized by comprising a staged combustion thrust chamber (6), an igniter (3), a turbopump, a starter (1), related valves and pipelines; the staged combustion thrust chamber (6) adopts a two-stage combustion mode and sequentially comprises a thrust chamber upper cavity (6-1), a gas-liquid injector (6-2), a thrust chamber lower cavity (6-3) and a tail nozzle (6-4) from top to bottom, wherein, part of oxidant and all fuel are organized to be burnt in a rich way in the upper cavity (6-1) of the thrust chamber, the gas-liquid injector (6-2) is provided with a rich gas injection unit and an oxidant injection unit, the rich gas generated in the upper cavity (6-1) of the thrust chamber and most of the oxidant after the oxidant pump enter the lower cavity (6-3) of the thrust chamber through the gas-liquid injector (6-2) to be burnt in a complementary way, the generated gas is exhausted out of the thrust chamber after being accelerated and expanded through the tail nozzle, the periphery of the thrust chamber is provided with a fuel cooling channel (6-5) of the thrust chamber, and the outlet is communicated with the upper cavity (6-1) of the thrust chamber; the external fuel is connected with the inlet of the thrust chamber fuel cooling channel (6-5) through a fuel pump (8) and a fuel valve (9); the external oxidant is respectively communicated with the upper cavity (6-1) of the thrust chamber and the gas-liquid injector (6-2) through an oxidant pump (7), an oxygen auxiliary valve (4) and an oxygen main valve (5); the turbine pump comprises a turbine (2), a fuel pump (8) and an oxidant pump (7) which are coaxially connected in series with the turbine, and a starter (1) is connected with a collector of the turbine; the turbine of the turbine pump is driven by rich fuel gas directly led out from an upper cavity (6-1) of the staged combustion thrust chamber, and the rich fuel gas is directly discharged to the outside after driving the turbine.
2. The open staged combustion bleed cycle liquid rocket engine of claim 1 wherein: the starter comprises a gunpowder starter or a high-pressure gas cylinder starter.
3. The open staged combustion bleed cycle liquid rocket engine of claim 1 wherein: and a thrust chamber oxidant cooling channel (6-6) is arranged outside the thrust chamber fuel cooling channel (6-5), and the external oxidant is communicated with the gas-liquid injector (6-2) through an oxidant pump (7), an oxygen main valve (5) and the thrust chamber oxidant cooling channel (6-6).
4. The open staged combustion bleed cycle liquid rocket engine of claim 1 wherein: the external oxidant is communicated with a main valve (12) through an oxidant pump (7), is directly communicated with the upper cavity (6-1) of the thrust chamber, and is communicated with a gas-liquid injector (6-2) through a main oxygen valve (5).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111251559.8A CN114060170B (en) | 2021-10-22 | 2021-10-22 | Open staged combustion air-extracting circulation liquid rocket engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111251559.8A CN114060170B (en) | 2021-10-22 | 2021-10-22 | Open staged combustion air-extracting circulation liquid rocket engine |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN114060170A true CN114060170A (en) | 2022-02-18 |
| CN114060170B CN114060170B (en) | 2023-05-23 |
Family
ID=80235759
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202111251559.8A Active CN114060170B (en) | 2021-10-22 | 2021-10-22 | Open staged combustion air-extracting circulation liquid rocket engine |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN114060170B (en) |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114718768A (en) * | 2022-04-02 | 2022-07-08 | 西安航天动力试验技术研究所 | High-temperature high-pressure high-speed large-flow-density oxygen-enriched gas treatment system and method |
| CN115355106A (en) * | 2022-08-24 | 2022-11-18 | 深圳驭龙航天科技有限公司 | Liquid rocket engine with combustion chamber for air extraction and circulation |
| CN116044610A (en) * | 2022-12-29 | 2023-05-02 | 北京航天动力研究所 | Double-expansion circulation liquid rocket engine system |
| WO2023221250A1 (en) * | 2022-05-18 | 2023-11-23 | 卢驭龙 | Coaxial liquid-propellant rocket engine using full-flow staged-combustion cycle |
| CN117329025A (en) * | 2023-12-01 | 2024-01-02 | 陕西天回航天技术有限公司 | Turbine exhaust stamping and pushing combined cycle engine and aerospace vehicle |
| CN117345471A (en) * | 2023-12-04 | 2024-01-05 | 陕西天回航天技术有限公司 | Self-adaptive height compensation spray pipe, engine and aerospace vehicle |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20040177603A1 (en) * | 2003-03-12 | 2004-09-16 | Aerojet-General Corporation | Expander cycle rocket engine with staged combustion and heat exchange |
| RU2531833C1 (en) * | 2013-07-17 | 2014-10-27 | Николай Борисович Болотин | Liquid propellant rocket engine |
| CN104919167A (en) * | 2013-02-19 | 2015-09-16 | 三菱重工业株式会社 | Rocket engine, rocket, and method for starting rocket engine |
| CN108412637A (en) * | 2018-03-16 | 2018-08-17 | 北京航天动力研究所 | A kind of novel hydrogen-oxygen rocket engine system |
| CN111622864A (en) * | 2020-06-03 | 2020-09-04 | 西北工业大学 | Semi-open type oxygen-enriched afterburning cycle engine |
| WO2021091613A2 (en) * | 2019-09-06 | 2021-05-14 | Special Aerospace Services, LLC | Staged combustion liquid rocket engine cycle with the turbopump unit and preburner integrated into the structure of the combustion chamber |
-
2021
- 2021-10-22 CN CN202111251559.8A patent/CN114060170B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20040177603A1 (en) * | 2003-03-12 | 2004-09-16 | Aerojet-General Corporation | Expander cycle rocket engine with staged combustion and heat exchange |
| CN104919167A (en) * | 2013-02-19 | 2015-09-16 | 三菱重工业株式会社 | Rocket engine, rocket, and method for starting rocket engine |
| RU2531833C1 (en) * | 2013-07-17 | 2014-10-27 | Николай Борисович Болотин | Liquid propellant rocket engine |
| CN108412637A (en) * | 2018-03-16 | 2018-08-17 | 北京航天动力研究所 | A kind of novel hydrogen-oxygen rocket engine system |
| WO2021091613A2 (en) * | 2019-09-06 | 2021-05-14 | Special Aerospace Services, LLC | Staged combustion liquid rocket engine cycle with the turbopump unit and preburner integrated into the structure of the combustion chamber |
| CN111622864A (en) * | 2020-06-03 | 2020-09-04 | 西北工业大学 | Semi-open type oxygen-enriched afterburning cycle engine |
Cited By (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114718768A (en) * | 2022-04-02 | 2022-07-08 | 西安航天动力试验技术研究所 | High-temperature high-pressure high-speed large-flow-density oxygen-enriched gas treatment system and method |
| CN114718768B (en) * | 2022-04-02 | 2024-01-12 | 西安航天动力试验技术研究所 | High-temperature high-pressure high-speed high-flow density oxygen-enriched gas treatment system and method |
| WO2023221250A1 (en) * | 2022-05-18 | 2023-11-23 | 卢驭龙 | Coaxial liquid-propellant rocket engine using full-flow staged-combustion cycle |
| CN115355106A (en) * | 2022-08-24 | 2022-11-18 | 深圳驭龙航天科技有限公司 | Liquid rocket engine with combustion chamber for air extraction and circulation |
| CN116044610A (en) * | 2022-12-29 | 2023-05-02 | 北京航天动力研究所 | Double-expansion circulation liquid rocket engine system |
| CN116044610B (en) * | 2022-12-29 | 2024-04-09 | 北京航天动力研究所 | Double-expansion circulation liquid rocket engine system |
| CN117329025A (en) * | 2023-12-01 | 2024-01-02 | 陕西天回航天技术有限公司 | Turbine exhaust stamping and pushing combined cycle engine and aerospace vehicle |
| CN117329025B (en) * | 2023-12-01 | 2024-02-23 | 陕西天回航天技术有限公司 | Turbine exhaust stamping and pushing combined cycle engine and aerospace vehicle |
| CN117345471A (en) * | 2023-12-04 | 2024-01-05 | 陕西天回航天技术有限公司 | Self-adaptive height compensation spray pipe, engine and aerospace vehicle |
| CN117345471B (en) * | 2023-12-04 | 2024-02-09 | 陕西天回航天技术有限公司 | Self-adaptive height compensation spray pipe, engine and aerospace vehicle |
Also Published As
| Publication number | Publication date |
|---|---|
| CN114060170B (en) | 2023-05-23 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN114060170B (en) | Open staged combustion air-extracting circulation liquid rocket engine | |
| CN113294264B (en) | Double-component variable-thrust rotary detonation rocket engine based on pintle injector | |
| CN111622864B (en) | Semi-open type oxygen-enriched afterburning cycle engine | |
| US8893467B2 (en) | Direct injection of a discrete quantity of fuel into channels of a wave rotor engine | |
| US8683780B2 (en) | Gas turbine engine and pulse detonation combustion system | |
| CN211082076U (en) | Pump pressure type engine started by blowing | |
| CN111502864B (en) | Open-cycle liquid oxygen kerosene engine system and use method thereof | |
| US6918243B2 (en) | Bi-propellant injector with flame-holding zone igniter | |
| US4901525A (en) | Booster-sustainer rocket engine and method | |
| RU2386844C1 (en) | Three-component liquid-propellant rocket engine and method of its operation | |
| CN113958425B (en) | Open-cycle liquid oxygen kerosene engine system and using method thereof | |
| RU2334916C1 (en) | Gas-dynamic igniter | |
| EP4030046A1 (en) | Multi-time ignition starting apparatus for a rocket engine, and rocket engine having same | |
| CN114526150B (en) | Jet ignition oxyhydrogen internal combustion engine based on precombustion chamber and control method | |
| CN116122989A (en) | RBCC combustion chamber with two-stage rocket layout and combustion organization method | |
| CN112360647A (en) | Multiple starting system of liquid rocket engine and starting control method thereof | |
| CN114352437A (en) | Solid fuel stamping combined engine suitable for wide Mach number flight | |
| CN107143432B (en) | High-piezoelectricity plasma gas relay couples spark knock engine before a kind of detonation wave | |
| RU2789943C1 (en) | Liquid rocket engine with accessor | |
| CN116044610B (en) | Double-expansion circulation liquid rocket engine system | |
| CN113530714B (en) | Pumping pressure type engine starting ignition method and system based on hydrogen peroxide | |
| CN112761821B (en) | Gas cylinder starting system and liquid rocket engine | |
| CN114320659B (en) | Low-cost closed pulse detonation wind source engine | |
| CN113123883B (en) | Turbine engine and self-starting method thereof | |
| CN112628023B (en) | Rocket engine ignition system and using method |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |